1. Field of the Invention
The present invention relates generally to the field of supersonic aeronautics, and relates more specifically to a novel supersonic or hypersonic body section design incorporating one or more hollow channels in the body section that allow freestream air to flow through the channel(s) resulting in reduced drag and sonic boom.
2. Description of Related Art
A variety of supersonic and hypersonic vehicles are currently being studied for commercial and military applications. However, existing supersonic and hypersonic technology suffers from several drawbacks. For example, aerodynamic drag is substantial at supersonic and hypersonic speeds, and adds significantly to fuel expenses. Also, noise due to sonic booms has been found objectionable, in particular, by those living under flight paths and near airports.
The advantages of supersonic and hypersonic transport, however, justify continued research and development directed to overcoming these drawbacks. For example, the high speed civil transport (HSCT) aircraft is designed to cruise at a free stream Mach number of approximately M∞=2.4 and seeks to overcome the economic and environmental barriers that have limited the success of previous supersonic commercial concepts. Other supersonic flight vehicles of significant interest include single-stage-to-orbit (SSTO) and multi-stage launch vehicles, tactical and strategic hypersonic and supersonic missiles, hypersonic cruise aircraft and planetary entry vehicles. These vehicles are similar in that their range, payload mass fractions and economic feasibility are extremely sensitive to aerodynamic drag.
A discussion of the effects of drag reduction on such supersonic vehicles is given by Bushnell, D., Supersonic Aircraft Drag Reduction, AIAA Paper 90-1596, 1990. If the lift-to-drag ratio (L/D) of the HSCT at cruise is increased by just 10% there would be a significant positive impact on the economy and success of that vehicle. Proposed hypersonic vehicles such as the National Aerospace Plane and X-30 have not advanced due, in part, to diminishing projected payload margins and concerns regarding airbreathing engine capabilities. As pointed out in Bushnell, drag reductions allow lower fuel requirements and can lead to reduced operating costs, as well as reduced sonic boom and noise effects. Reviews of supersonic drag reduction techniques also are given by Hefner, J. N., and Bushnell, D. M., An Overview of Concepts for Aircraft Drag Reduction, AGARD Rep. 654, June, 1977, pp. 1-1 to 1-30; and by Jones, R. T., Aerodynamic Design for Supersonic Speeds, Adv. in Aero. Science, Vol. 1, 1959, pp. 34-52
A schematic of a typical drag breakdown, taken from Kuchemann, D., The Aerodynamic Design of Aircraft, Pergamon Press, Ltd., 1978, for supersonic vehicles (M∞=2.4) is shown. The vehicle drag coefficient is shown as a function of the product of xcex2={square root over (M∞2xe2x88x921)} and the semispan to length ratio, s/l. It should be noted that vehicles with s/l lower than approximately 0.2 (i.e. vehicles whose wings are extremely swept) are impractical due to excessive required runway lengths. Note also that s/xcex2l=1 corresponds to the case in which the Mach number normal to the wing leading edge is sonic. The drag on supersonic vehicles can be classified into three different categories: (1) skin friction drag, (2) drag due to lift, and (3) zero-lift bluntness (thickness-wave) drag. Skin friction drag is due to fluid viscosity and is a function of the total wetted surface area of the vehicle. Drag due to lift consists of induced drag and the component of wave drag which is a function of the inclination of the vehicle surfaces with respect to the freestream direction at a non-zero lift orientation. Finally, the zero-lift bluntness drag is the wave drag due to the thickness and bluntness of the leading and trailing edges of the vehicle in a zero lift orientation. The zero-lift bluntness drag (i.e. thickness-wave drag) increases rapidly with freestream Mach number and can be responsible for well over ⅓ of the total vehicle drag.
Linearized supersonic theory indicates that for an airfoil of a given thickness, the shape which gives minimum zero-lift bluntness drag is the sharp diamond airfoil. However, very sharp leading edges are not practical for a number of reasons: (1) very sharp leading edges are difficult and expensive to manufacture; (2) some blunting is required for structural strength, (3) the flow over wings with sharp leading edges is very susceptible to separation even at low angles of attack and flight speeds; and (4) the heat transfer to sharp leading edges at high supersonic Mach numbers is severe.
For hypersonic vehicles, heat transfer considerations often dictate the design of the nose and the leading edges. The heat transfer to such vehicles is most severe at stagnation points which occur on the leading edges and nose of the vehicle. Theoretical and numerical predictions of stagnation point heating have been developed by Fay, J., and Riddell, F., Theory of Stagnation Point Heat Transfer in Dissociated Air, J. Aero. Sci., Vol. 25, pp. 73-85, February, 1958, and are also described by Anderson, J. D., Hypersonic and High Temperature Gas Dynamics, McGraw-Hill, Inc., New York, 1989. Kemp and Riddell have developed an accurate semi-empirical relation for stagnation point heat transfer. Kemp, N., and Riddell, F., Heat Transfer to Satellite Vehicles Re-entering the Atmosphere, Jet Propulsion, pp. 132-137, February, 1957. Theoretical formulations, experimental data, and semi-empirical formulas all agree in the fact that stagnation point heat transfer is inversely proportional to the square root of the nose or leading edge radius, i.e.,       q    stag    ∝      1                  r        n            
Thus, sharp leading edges (i.e., rn=0) experience extreme heat transfer, which may melt or otherwise damage the component. Therefore, the nose, leading edges of wings, tails, fins, struts, cowls, and other appendages of supersonic and hypersonic vehicles are blunted so that the heat transfer and structural loads will be manageable. With leading edge blunting, the simple diamond airfoil is modified. However, much of the wave drag experienced by these vehicles is due to leading edge blunting. Wave drag is responsible for approximately one-third of the total drag experienced by aircraft, atmospheric entry vehicles, missiles, and other vehicles in supersonic and hypersonic flight.
For vehicles which cruise at low supersonic Mach numbers, heat transfer considerations do not dictate the design of the wing leading edges. At subsonic, off-design conditions, such as takeoff, landing, climb and maneuvering flight, blunted leading edges are desirable so that flow separation is prevented. However, a blunted leading edge results in higher drag at supersonic cruise conditions relative to a wing with a sharp leading edge. Therefore, it has been found desirable that airfoils utilized in such applications be significantly blunted during subsonic maneuvering phases of flight, but perform more like sharp leading-edge airfoils at supersonic cruise conditions. The provision of such an airfoil is one object of the present invention.
Sonic boom has also been found to limit the application of supersonic and hypersonic transport. To date, the most successful strategies utilized for minimization of sonic boom stem from evaluation of the Whitman F-Function. Whitman, G. B., The Flow Pattern of a Supersonic Projectile, Communications on Pure and Applied Mathematics, Vol. V., 1952, pp. 301-348. See also Carlson, H. W., and Maglieri, D. J., Review of Sonic Boom Generation Theory and Prediction Methods, Journal of Acoustical Society of America, Vol. 51, 1972, pp. 675-685; and Middleton, W. D., and Carlson, H. W., A Numerical Method for Calculating Near-Field Sonic-Boom Pressure Signatures, NASA TN D-3082, 1965. The F-Function is based on the cross-sectional area distribution and the lift distribution of the generating vehicle. Modified linearized theory and geometric acoustics are used to determine ground pressure and signatures from the F-Function. Studies by Seebass and George, and Darden have demonstrated that mid-field and far-field over-pressures are minimized when a perfectly blunt (i.e. flat face) nose is utilized on the vehicle. Seebass, R. and George, A. R., Sonic Boom Minimization, Journal of the Acoustical Society of America, Vol. 51, No. 2, Part 3, 1972, pp. 686-694; Darden, C. M., Sonic Boom Theory: Its Status in Prediction and Minimization, Journal of Aircraft, Vol. 14, No. 6, 1977, pp. 569-576.
When a blunt nose is utilized, the subsequent compressions are weak. However, for sharp-nosed vehicles, subsequent compressions around the vehicle tend to be strong and coalesce at mid-field and far-field into a much stronger over-pressure relative to a blunt nose geometry. Significantly blunted supersonic cruise vehicles have not been utilized, however, because the wave drag of such vehicles becomes unreasonably high. The relationship between nose bluntness and boom characteristics has been validated experimental studies. Carlson, H. W., Mack, R. J., Morris, O. A., A Wind Tunnel Investigation of the Effect of Body Shape on Sonic-Boom Pressure Distributions, NASA TN D-3106, November 1965, and Mack, R. J. and Darden, C. M., Wind Tunnel Investigation of the Validity of a Sonic-Boom-Minimization Concept, NASA TP-1421, 1979. The studies by Darden of vehicles with a relaxed level of nose bluntness clearly show that a minimum boom must be paid for by higher drag penalties.
Optimization studies have sought to find a palatable compromise for this xe2x80x9clow-boom, high-drag paradoxxe2x80x9d but breakthroughs in design of aerodynamically efficient, low boom vehicles require consideration of innovative approaches in vehicle design. Various approaches are described, for example, by: Haglund, G. T., HSCT Designs for Reduced Sonic Boom, AIAA Paper 91-3103, AIAA Aircraft Design Systems and Operations Meeting, September 1991; Mack, R. J. Needleman, K. E., A Methodology for Designing Aircraft to Low Sonic Boom Constraints, NASA TM 4246; Yoshida, K., and Tokuyama, A., Improving the Lift to Drag Characteristics of Low Boom Configuration, AIAA Paper 92-4218, August, 1992; Yoshida, K., Experimental and Numerical Study for Aerodynamics of Low Boom Configuration, AIAA Paper 94-0052, Aerospace Sciences Meeting, January 1994; and Darden, C. M., Sonic Boom Theory: Its Status in Prediction and Minimization, Journal of Aircraft, Vol. 14, No. 6, 1977, pp. 569-576.
It is known to provide a reverse flow airfoil in which air is actively pulled into the trailing edge of an airfoil through a duct and ejected forward out of the leading edge of the airfoil. Kucheman, D., Some Aerodynamic Properties of a New Type of Aerofoil with Reversed Flow Through an Internal Duct, RAE TN Aero 2297, 1954. Several studies have been performed for supersonic flow in which either air or some other gas is forcibly ejected forward out of the blunt nose of axisymmetric bodies. Kucheman, D., Some Aerodynamic Properties of a New Type of Aerofoil with Reversed Flow Through an Internal Duct, RAE TN Aero 2297, 1954; Finley, P. J., A Preliminary Investigation of the Steadiness of a Free Stagnation Point, J. Dept. of Engineering, University of Malaya, Vol. 4, 8, 1965; Finley, P. J., The Flow of a Jet from a Body Opposing a Supersonic Free Stream, Journal of Fluid Mechanics, Vol. 26, 337, 1966; McMahon, H. M., An Experimental Study of the Effect of Mass Injection at the Stagnation Point of a Blunt Body, GALCIT Memo 42, 1958; Lam, S. H., Interaction of a Two Dimensional Inviscid Incompressible Jet Facing a Hypersonic Stream, AFOSR R 447, 1959; Eminton, E., Orifice Shapes for Ejecting Gas at the Nose of a Body in Two Dimensional Flow, RAE TN Aero 2711, 1960; Warren, C. H. E., An Experimental Investigation of the Effect of Ejecting a Coolant Gas at the Nose of a Blunt Body, GALCIT Memo 47, 1958, and Journal of Fluid Mechanics, Vol. 8, 400, 1960; and Reid, J., and Tucker, L. M., Forward Ejection From Swept and Unswept Leading Edges, RAE TR 68095, 1968. This technique causes the existence of a stagnation point in front of the duct and is a means of providing active cooling at the nose. However this technique suffers the disadvantage that it requires that significant power be supplied from the engines to pump the flow against a supersonic freestream. By contrast, in the present invention, only passive flow of air through the channel is necessary.
Thus, it can be seen that a need yet exists for a supersonic and/or hypersonic body configuration which minimizes the body""s drag without adversely affecting its lift characteristics.
A need further exists for a method of reducing drag on airfoils and other bodies in supersonic or hypersonic flow, without adversely affecting their lift characteristics.
Still a further need exists for a method of reducing sonic boom associated with a supersonic or hypersonic body while reducing the total drag of the body as compared to conventional blunted geometries.
A need also exists for a supersonic body configuration which presents an overall flow structure similar to that of conventional, low sonic boom, blunted geometry body configurations, but with reduced drag.
It is to the provision of a method and apparatus meeting these and other needs that the present invention is primarily directed.
Briefly described, in preferred form, the present invention comprises a method of reducing the drag and the sonic boom generated by a supersonic body. The present invention also comprises a supersonic body configuration having an improved lift/drag ratio and reduced sonic boom.
Instead of significant blunting of the leading edge, the present invention provides a hollow channel in the vehicle nose and in airfoil sections that make up the wings, tails, fins, struts, cowls, and other appendages of supersonic and hypersonic vehicles. The channel preferably begins at or adjacent the leading edge of the component, with the freestream air flowing through the channel. The present invention thereby provides significantly reduced wave drag and total drag, including skin friction drag, for a wide range of flight conditions. With a channel of a preferred size, a normal shock exists in front of the airfoil""s leading edge, and the surrounding flow is similar to that in front of a blunted airfoil. The heat transfer at the leading edge for this channeled airfoil is also similar to that for a blunted airfoil; however, the wave drag is significantly lower.
The present invention primarily addresses reduction of the zero-lift bluntness drag as a means of reducing the total drag. A reduction of this component of drag and in the total drag can result in increased vehicle range, increased speed, improved fuel efficiency, increased lift/drag ratio, and enhanced performance. An investigation of the drag reduction provided by the present invention for supersonic airfoils and wings has been conducted, and is reported below. A wide range of geometric parameters and supersonic flight conditions are considered, so that the aerothermodynamic performance of airfoils employing the present concept are generally characterized.
Accordingly, it is an object of the present invention to provide a method and apparatus for the reduction of drag and the enhancement of the lift-to-drag ratio of a supersonic body.
It is another object of the present invention to provide a method and apparatus for reducing sonic boom associated with a supersonic body, without adversely affecting the drag on the body.
These and other objects, features and advantages of the present invention will become more apparent upon reading the following specification in conjunction with the accompanying drawing figures.